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Delft University of Technology

A recovery system for the key components of the first stage of a heavy launch vehicle

Dek, Casper; Overkamp, Jean Luc; Toeter, Akke; Hoppenbrouwer, Tom; Slimmens, Jasper; van Zijl, Job;

Areso Rossi, Pietro; Machado, Ricardo; Hereijgers, Sjef; Kilic, Veli

DOI

10.1016/j.ast.2020.105778

Publication date

2020

Document Version

Final published version

Published in

Aerospace Science and Technology

Citation (APA)

Dek, C., Overkamp, J. L., Toeter, A., Hoppenbrouwer, T., Slimmens, J., van Zijl, J., Areso Rossi, P.,

Machado, R., Hereijgers, S., Kilic, V., & Naeije, M. (2020). A recovery system for the key components of the

first stage of a heavy launch vehicle. Aerospace Science and Technology, 100, [105778].

https://doi.org/10.1016/j.ast.2020.105778

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Contents lists available atScienceDirect

Aerospace

Science

and

Technology

www.elsevier.com/locate/aescte

A

recovery

system

for

the

key

components

of

the

first

stage

of

a

heavy

launch

vehicle

Casper Dek,

Jean-Luc Overkamp,

Akke Toeter,

Tom Hoppenbrouwer,

Jasper Slimmens,

Job van Zijl,

Pietro Areso Rossi,

Ricardo Machado,

Sjef Hereijgers,

Veli Kilic,

Marc Naeije

TUDelft,FacultyofAerospaceEngineering,Kluyverweg1,2629HSDelft,theNetherlands

a

r

t

i

c

l

e

i

n

f

o

a

b

s

t

r

a

c

t

Articlehistory:

Received2October2019

Receivedinrevisedform7February2020 Accepted10February2020

Availableonline17February2020 CommunicatedbyMehdiGhoreysh

Keywords:

Ablator

Controllableparafoil Firststagerecovery Inflatableheatshield Mid-airrecovery Reusability

Inrecentyears,thespacemarkethasbeenpushingtowardsdecreasingcostsoflaunchingspacecraftby reusingpartsofthelaunchers.Thepurposeofthisarticleistopresentafeasibilitystudyofarecovery systemfor theengine and engine frameofan existing,expendableheavy launchvehicle and present recommendationsforfurtherresearch.

The concept developed is verifiedbased onthe Ariane 6.The recoveryof the VulcainAftBay(VuAB)

isinitialised byseparationfromthe firststageat157.7 kmaltitudewhiletravellingat6930 m/s.The studyinvestigatesaninflatableaeroshellforprotectionand decelerationduringre-entry,afterwhichit isproposedtofurtherdeceleratetheVuABusingdrogueparachutes.Thefinalpartoftheconceptentails retrievaloftheVuABbyahelicopterinmid-air.To enableacontrolledglidingflightduringretrievala parafoilisproposed.Atlaunch,therecoverysystemwillweigh2789kgwithapayloadpenaltyof720 kg. Thesystemcan beintegratedintothe existingdesign ofthelauncher and willnot interferewith nominaloperationsofthelauncher.Implementingtherecoverysystemcanreducethecostperlaunchof anAriane6by15%

©2020TheAuthors.PublishedbyElsevierMassonSAS.ThisisanopenaccessarticleundertheCC BY-NC-NDlicense(http://creativecommons.org/licenses/by-nc-nd/4.0/).

1. Introduction

Inrecentyears themarket forspacelaunch vehicles hasseen significantchange.Privatecompanieshavebeenincreasingly inno-vative,pushing towards significant cost reductionsandincreased economiesofscale.Asofrecently,especiallyre-usabilityhasbeen ofgreat interest. Lowerstages oflaunchers, due tothe relatively lowseparationfromthelauncher,arethefirsthurdletotake.

To assess the potential of recovery, a technology review has been conducted to evaluate the potential of new recovery sys-temmethods.The mostpromising concepts fromthisreview are utilisedaspartofa feasibilitystudy to investigatethepossibility ofintegratingtheseinnovativerecoverysystemelementsinan ex-istinglauncherdesign.Thegoalistoinducelowpayloadpenalties, low redesign investmentrequirements andachieve a decreasein costoveralaunchcampaignofseverallaunchesinordertoenable existinglauncherstobepartlyreusable.

Thisworkis theresultofthe BScdesignsynthesisexerciseat theTU Delft

FacultyofAerospaceEngineering(2018).

*

Correspondingauthor.

E-mailaddress:m.c.naeije@tudelft.nl(M. Naeije).

This paper provides a trade-off of promising new technology fromthetechnologyreview,followedby afeasibilitystudy ofthe mostpromisingconcept.Themissionprofileisdiscussed,followed by adescriptionofthedifferentphasesofflight andrecovery. Fi-nally, the integration ofthe recoverysystem andthe launcher is discussed,leadingup toadiscussiononthefeasibilityofthefirst stage heavylaunchvehicle recoverysystem, followedby the con-clusionandfuturerecommendations.

This paper is based on the thesis report of 10 Bachelor Aerospace students at Delft University of Technology (TU Delft). The research hasbeen conductedusingthe Ariane 6launcher as abasis. Technicaldataforthispurely academicexercisewas pro-videdby TUDelft, through cooperationwithAirbus Defence and SpaceNetherlands(the ’client’).TheAriane6 firststagewas cho-sen, since its high separation velocity and altitude would make thisachallengingengineeringexercise.Inquiryintothefullreport canbemadetothecorrespondingauthor.

2. Methodology

The overall methodology and steps taken inthis research are presentedin thissection. Forthisresearch, severalrecovery con-cepts have been generated andanalysed on a top-level manner. The concepts were generated based on similar research and

ex-https://doi.org/10.1016/j.ast.2020.105778

1270-9638/©2020TheAuthors.PublishedbyElsevierMassonSAS.ThisisanopenaccessarticleundertheCCBY-NC-NDlicense (http://creativecommons.org/licenses/by-nc-nd/4.0/).

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2 C. Dek et al. / Aerospace Science and Technology 100 (2020) 105778

isting solutions for recovery methods of objects from space and high-altitudeflight.Afterconceptgeneration,theseconceptswere analysed in a comparative manner, using both quantitative and qualitative measures. The quantitative measures include the in-ducedweightofthetotalsystem, aswellastheinduced payload penalty of this extra weight. The sizing of the systems for this analysis was performed using similar methods as introduced in latersectionsofthisarticle.Also,theestimateddevelopment and per-itemcostofthe conceptwere takenintoaccount.Qualitative measures include the ability of the concept to safely return the objecttoEarth, developmenttimeanddevelopmentrisksandthe integrationpossibilitiesinexistinglaunchers.Forthelast threeof these, the goal in mindwas to be able to launch the newly de-signedsystemonanexistinglauncherwithin5yearstime.

Basedonthistrade-offa conceptwaschosen, whichwasthen furtheranalysed. Thisanalysis was performedby firstidentifying the full mission profile to be executed by the recovery system. Basedonthismissionprofile,thesizing ofallcomponentsofthe recoverysystemwasperformed.Thesizingofthecomponentsand determining the overall mission profile was an iterative process considering that the weight of the components has a direct re-lationtothevelocitiesanddistancescoveredinallmissionphases and vice-versa. The methods used to perform the sizing of the componentsandtoassesstheoverallfeasibilityoftheconceptare introduced in separate sections accompanied by the results and outcomesofthesemethods.

3. Conceptgenerationandtrade-off

The mission of the recovery system is to recover the engine bay of the first stage of the Ariane 6 and to allow it to be re-furbishedandreusedforsubsequentlaunches.Recoveryshallstart afterseparation ofthe first stage at157.7 km altitudeabove the AtlanticOcean,travellingat6930m/svelocity [18,6].Followinga literature study and evaluation ofcomparable missions, multiple conceptshavebeendeveloped.Usingdesignoptiontreelogic, fea-sibleoptionsareselected.Thesedesignoptionsarethencompared in a trade-off that evaluated the TechnologyReadinessLevel(TRL),

themissionrisk, thepredictedpayloadpenalty,thecosts,andthe developmentriskofeachconcept.

For comparison, the currently best performing and feasible reusable systems from the aerospace sector are the reusable booster stagesemployed by SpaceXon their Falcon9 andFalcon Heavylaunchers[2].SpaceXoperatesafullyreusablefirststageon theirFalcon9andFalconHeavylaunchers,whichare steeredand decelerated through a combination of thrust from the reignited and gimballed engines and control from grid fins. Additionally, AirbusDefence andSpaceproposed a concept, Adeline,that fea-turedfixed wings, actuatedcontrol surfaces andelectricengines, mountedontheenginebay[1].Thiswillallowforaconventional fixedwinglanding onarunway,eitheronlandoronaboat. Fur-thermore, two concepts utilising a combination of an inflatable heat shieldandan autonomousguided parafoilhave been evalu-ated.Thefirstperformsasoftlandinginalargeboat-mountednet, whilethesecondisretrievedthroughaMid-AirRecovery(MAR)

us-ingahelicopter.

In the trade-off, the concept similar to Adeline hasbeen dis-carded due to a disproportional high development risk as their design willinterfere withthe existing launcherdesign. Addition-ally,thehighenvisionedcostsandhighpayloadpenaltymakethis infeasible.Similarly,aSpaceX-typeverticalthrustpoweredlanding hasbeendiscarded.Thisisbecausethesignificantlylargervelocity andaltitudeoftheAriane6atthefirststageseparationin compar-isontotheFalcon9meanseitherexcessiveamountsofpropellant fordeceleration beforere-entryarerequiredorare-designofthe missionprofile isneeded. The analysed concepts andreasons for

Table 1

Analysedconceptsandreasonsforrejection.

Concept Mainreasonforconceptrejection Horizontal fixed wing landing Largeinterferencewithdesignand

operationsofexistinglaunchers Vertical thrust-powered landing Highseparationaltitudeandspeed,

resultinginlargeamountofpropellant andhighinducedpayloadpenalty Parafoil - net landing Highlysensitivetoerrorsinlanding

position

Mid-air recovery –(selectedconcept)

Fig. 1. Mission profile for the mid-air recovery concept.

Table 2

MissionprofiledetailsasvisualizedinFig.1,’Distance’isgrounddistancewith re-specttoseparation.

Mission event Alt. [km] Time [s] Distance [km] 1. Separation 157.7 0 – 2. Re-Entry 100 240 1634

3. Drogue parachute deployment 8.1 724 2844 4. Parafoil deployment 7.3 734 2844 5. Start alignment phase 4 1248 2854 6. Start catch window 1.2 1759 2852

rejectioncanbe foundinTable1.Acontrolled landingina boat-mounted net was discarded due to the slow manoeuvrability of boats large enough forsuch a catch. Finally,a MAR concept was selected, owingto itsrelatively low massandthe abilityto inte-gratethesystemwithminimumredesignoftheoriginallauncher.

4. Missionprofile

With the MAR concept selected for the feasibility study, the missionprofileisdeterminedandvisualizedinFig.1.Table2gives the details about the different mission phases. The Ariane 6 is scheduled to launch from Guiana Space Center. After liftoff, first and second stage separation is performed. Withno flown trajec-tories available from the Ariane 6, dataon the Ariane 5 is used fortheseparationaltitudeandaccompanyinglaunchervelocity,as the Ariane6 isexpectedtohave asimilar trajectory. The separa-tionaltitudeandlaunchervelocityarefixedat157.7kmand6930 m/srespectively[6]. Thiswasconfirmedby ourcontactatAirbus DefenceandSpaceNetherlands(2018).Untilthispoint,nothingis changedfromtheoriginallaunchsequenceoftheAriane6inorder nottointerferewithnormaloperations.Next,theVuABseparation from thefirst stage occurs. An inflatable aeroshellwill deploy in

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Fig. 2. Velocity of the VUAB after the space flight phase.

space,followedbyatmosphericre-entryatavelocity of7005m/s andflight path angle of-20◦. At 8 kmaltitude and a VuAB ve-locityof85 m/sa drogueparachutedeploys,followedby parafoil deployment at 27m/s. This occursat an altitude of 7.3 kmand depictsthestart of theguided parafoil descent.This parafoil en-sures a horizontal and vertical flight velocity of 18 m/s and4.5 m/s,respectively.From4kmaltitudeonward,theparafoiland he-licopter willbegin alignment procedures. At 1.2km the catching phase commencesandthe parafoil velocitiesare 15 m/sand3.9 m/shorizontalandverticalrespectively.In Fig.7itisshownthat thesevelocitiesallowforahelicoptertorecovertheVuABin mid-air.Afterthe VuABisrecovered,it islandedonarecoveryvessel whichshipsthe VuABto France inorderforit to be refurbished andreused.

5. Spaceflightphase

ThemissionoftheVuABrecoverysystemcommenceswiththe separationoftheVuABfromthefirststage.Thisseparationis per-formed with pyrotechnics and induces rotational motion on the VuAB[18].Thisrotationalmotionneedstobedamped,inorderto positiontheVuABforre-entry.

Thespaceflight phasestartsat157.7kmaltitude,endsat100 kmaltitudeand its durationis approximated at4minutes using a trajectory calculation. Forthis purpose,a space flight modelis developedthatmodelsandsimulatesaballistictrajectorywithan

AttitudeDeterminationandControlSystem(ADCS) for attitude con-trol. In this model, the angular momentum per unit mass, H, is used,whichisconstantthroughoutthespaceflightphaseand de-terminedusingEquation (1):

H

=

Vt

·

Re

=

V

· (

Re

+

h

)

(1)

where Vt is thetangential velocity, Re is theradius oftheEarth andh isthe altitudeofthe VuAB,whichisdefinedas100kmat thepointofre-entry[22].Usingthisequation,thevelocitiesatthe endofthespaceflight phasewe determined,asshowninFig.2. TheshapeoftheballistictrajectoryisdefinedusingEquation (2):

p

=

H

2

μ

(2)

assuming an elliptical shape of the trajectory and p being the semi-latusrectum definingthatshape,with

μ

beingthestandard gravitationalparameterforEarth. The final spaceflight trajectory ispredictedtobeasdepictedinFig.3.

Fig. 3. Predicted Space Flight Phase Trajectory of the VUAB.

Themodelisabletopinpointthere-entrylocationgiveninitial separationconditionsasstatedinthemissionprofile.Acontroller is usedto ensure theVuAB is inthe desirableorientation atthe startofthe re-entrymissionphase, andmakes pitchadjustments whenrequired.Thisallowstheattitudecontrolthrusters,including the requiredpropellant, tobe sized. Based on amaximum space flighttimeof4minutesandaslewrateof8.6deg/satseparation [18],theDST-13thrusterbyMOOG[28] isselected.Withathrust of 22N anda specific impulse of 300s, the requiredpropellant mass to fuel the 16 thrusters,four thrusters at each of the four sides,iscomputedat0.7kg.

6. Re-entry

At100kmaltitude,travellingat7005m/s,theatmospheric ef-fectsbecome non-negligible.The highvelocity atre-entryresults inhighdynamicpressuresandhighheatflux,suchthataThermal ProtectionSystem(TPS) isrequiredtoprotect theVuABfromthese aero-thermalloads.Due tothe limitedamountofvolume on the VuABarigidaeroshellcannotbeintegrated,sothereforean inflat-ableaeroshellisdesignedforthispurpose.Thisaeroshellisbased onexistingdesigns byNASA,namelytheHEART[10] andIRVE-III [8] concepts.

Aglidingre-entrymodelisdevelopedtosimulatethere-entry trajectory and conditions. The re-entry model uses a linearised gravitymodelforasphericalEarth,withtheInternationalStandard Atmosphere(ISA)andconsistsofalinearisationoftheCD

Mach

relationof[26] for M

>

2 foraninitialangleofattack

α

= −

20◦. Thestability oftheVuAB duringre-entrywithdeployed aeroshell isanalysedbycomparisonwiththeaeroshellin[19],whichis sta-bleat

α

= −

20◦.Toincludesmallvariationsintheangleofattack in the reentry model,the relation between CD and

α

of [19] is linearisedaround

α

= −

20◦.Theinitialconditionsofthere-entry aretakenfromtheendofthespaceflightphaseandintegratedto sizetheaeroshellinaniterativeprocess.Firstthenosecone diam-eterDi,aeroshelldiameterD0,spherical-coneangle

θ

,nose-radius

Rnose, toroid diameter Dt and number of stacked toroids N, are sizedanddeterminedusingEquation (3) [25,24,21]:

Dt

=

Do

Di

(

2N

1

)

sin

a

)

+

1

cos

a

)

(3) withaninitialvalueofN

=

7 inthefirstiteration.Furthermore,D0

isconstrainedtoaminimumdiameterof9.14mtoprevent inter-ferenceoftheincomingflowwiththeVuABunderanangleof at-tackof

α

= −

20◦.Subsequently,theheatfluxanddynamicpressure

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4 C. Dek et al. / Aerospace Science and Technology 100 (2020) 105778

Table 3

Aeroshellcomponentmassbreakdown.

Components % Mass [kg]

Adapter 22.1 311.4

Heatshield support stucture 5.1 71.9 Inflation mechanism 16.8 236.7 Separation mechanism 12.4 174.7

TPS 43.6 614.3

Total 100 1409

Fig. 4. Cross-section view of the aeroshell components.

aredetermined forlaminarcompressibleflow, usingEquation (4) and(5) [4]:

˙

qw

=

ρ

V3



1

.

83

·

10−8

Rnose

 

1

hw h0



(4) q

=

1 2

γ

psM 2 ∞ (5)

inwhichhw isthe enthalpy atthewall,h0 isthetotal enthalpy,

ρ

isthefreestreamdensity, V isthevelocity, ps isthestatic pressure and M is the free stream Mach number. Since the aeroshellcanbeconsideredabluntbody,themaximumheatflux is assumedto occur at the stagnationpoint of the aeroshell [4]. Themaximumheatfluxandmaximumdynamicpressureareused toselecttheTPSmaterial.Thisyields aconservativeTPSmaterial selectionsincetheheatfluxwillbelowerthanthemaximum heat-fluxoutsidethestagnationpoint.Next,N,D0,

θ

,Rnose,Dt andthe

selectedTPS materialproperties aresubstituted ina massmodel basedon amodified liftingHIAD massmodel [33],which results inTable3.

The final aeroshell design has a diameter of 8.8 m, uses the ablative SIRCA-flex asthe TPS material and has a mass of 1409 kg. Fig.4 shows the cross-section ofthe aeroshell in its inflated state,clearlyshowingtheinflatabletoroidstructure.Theaeroshell isdesignedtofitintotheenginebaybymeansofdeflationofthe toroidsandsubsequentfoldingofthedeflatedstructure,ascanbe seen in thecross-section in Fig. 13,in which the beige-coloured structurerepresentsthefoldedandpackagedaeroshell.

Inordertoverifytheoutputsofthedevelopedreentrymodel, thesamedesignparametersastheHEARTaeroshellareusedasan inputto themodel,afterwhich theoutputs, plottedinFig.5are compared with the measurements performed during the HEART testasdisplayedinFig.6[10].

Ascanbe seenwhencomparingFig.5withFig.6,the magni-tudeofthe outputs is comparable.Thedifference between maxi-mumheatratefoundatthesurfaceislikelyduetoadifferencein aeroshellgeometrysuchasthenose-coneradius.Thephenomenon ofhigherdynamicpressureandaccelerationfortheHEARTislikely causedbya higherkineticenergyatthestartofthe atmospheric entry,asthistestisbeingperformedafterseparationofthe Inter-nationalSpaceStation.

Inthenext 500seconds,theaeroshelldeceleratestheVuABto around85m/s,coveringagrounddistanceof1200km.A

simula-Fig. 5. Verification ofthere-entrymodelusingthesamedesignparametersusedfor theHigh-EnergyAtmosphericRe-entryTest[10].

Fig. 6. Measurements performedduringtheHigh-EnergyAtmosphericRe-entryTest [10].

tionoftheverticalre-entrytrajectoryisperformedusingthesizing data from the aeroshell, varying atmosphericdata and the vary-ingaerodynamicparametersthroughoutre-entry.Theaerodynamic parameters ofthe aeroshellare estimatedusing relationships be-tween angle of attack, mach number and the aerodynamic lift, moment and drag coefficients, as defined by [26] and [19]. The resultsofthissimulationcanbeseeninFig.7andFig.8.

Inreality,atmosphericconditionsarestochastic,henceaMonte Carlosimulationofthere-entryflightphaseisperformedto eval-uatetheaccuracyofare-entryusinganaeroshell.Thisisdoneby varying atmosphericconditionsandflight parameterssuch as an-gleofattackandtheaerodynamiccoefficientsoveraspanof1000 simulations, whichcan beseen inFig.9.The variation inthe in-putsduring thisMonteCarlosimulationaregiveninTable4.The decelerationisshowninFig.10,showingamaximumdeceleration ofalittleover6 gataround100scorrespondingtoanaltitudeof 60km(Fig.7).

Theaccuracyoftherecoverystronglydependsonatmospheric conditionsandexternalforcesandisstronglyinfluencedbyslight

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Fig. 7. Altitude of the VuAB versus time during the re-entry phase.

Fig. 8. Velocity versus time of the VuAB during the re-entry phase.

Fig. 9. Scatterplot of the Monte Carlo Simulations.

delays in separation. However, the Monte Carlo simulations do indicate that upon variation of atmosphericconditions, aeroshell surfacearea andangleof attack,the VuABwould notendits re-entryphaseoutsideoftherangeoftheretrievalhelicopter.Inthe worstcasescenario,there-entryphaseoftheVuABended15km fromtheexpectedlocation (seeFig.9),whichmeans thatthe re-trievalhelicopterwould stillbe capable ofreaching itwithin the durationoftheatmosphericflightphase.

Table 4

UncertaintiesasusedintheMonteCarloSimulation. Variable Random deviation (2σ) Aeroshell surface area 2%

Air temperature 5% Air pressure 5% Air density 5% Angle of attack 2.7 deg Sideslip angle 2.7 deg

Fig. 10. Deceleration oftheVuABversustimeforoneoftheatmosphericentry sim-ulationsperformedduringtheMonteCarloanalysis.

Theatmosphericentrydistributionshowsapredictedrangeof potential locations at which the atmospheric flight phase starts, thesizeofthisdistributionis26kmlongitudinallyand4km lat-erally.Thisrangeofuncertaintyinthere-entrypositionservesas an inputtothe requiredparafoilperformance andcontrol system sizing.

The aeroshell must be discarded before drogue parachute deployment. Due to volume and temperature constraints, the parachute and aeroshell deployment direction is oriented oppo-site to the nozzle. Therefore the VuAB needs to be flipped 180 degrees along the lateralaxis before parachute deployment. This discarding and rotation is performed by gradual deflation of the aeroshellwhichcausesapitchupmomentthat rotatestheVuAB, allowingtheaeroshelldiscardedbymeansofaspringmechanism.

7. Atmosphericflightandrecovery

In order to transition from the re-entry phase to the atmo-spheric flightphase,a drogueparachuteissized whichisableto deceleratethe VuABfurther. Thisdrogueparachute usesa forced ejectionbymeansofamortar,whichissizedat26kgbymeansof a comparativestudy into mortarsforparachutedeployment [16]. The parachute andparafoil deployment system is constrainedby amaximumallowed decelerationof4.5g[18], anda descent ve-locitybelow 5m/sto providethe helicopterwithsufficient time toreach andcapturethe VuAB.Thisprovides thelimitsinwhich theparachutehastooperate.Theparachuteissizedusingthe as-sumptionsthattheVuABwillexperiencea2seconddelaybetween discardingtheaeroshell andparachutedeployment,resultingina free-fall. Also, during deployment the VuAB is assumedto move in near-vertical direction and the drag coefficient is assumedto be similar tothe parachutesof theOrion [3]. Theparachute

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sys-6 C. Dek et al. / Aerospace Science and Technology 100 (2020) 105778

Fig. 11. The acceleration, altitude, velocity and g-forces throughout the deployment phase of the drogue parachute and parafoil.

Table 5

Parachutesystemmassesandtheirrespectivevolumes.

System Mass [kg] Volume [kg/m3]

Drogue parachute 47.00 0.065 Main parafoil 435.5 0.604 Secondary parafoil 75.67 0.105 Mortar 26.00 0.036 Control system 32.66 0.045 Margin 123.4 0.171 Total 740.1 1.028

temwasdesignedandsizedusingtheOrionmissionasareference [15] andusingamodelconsistingofsimpleliftanddragequations asafirstapproximation.Themassandpackingvolumeofall com-ponentsofthesystemcanbeseeninTable5.

The drogue parachute decelerates the VuAB to a velocity of 27 m/s atan altitude of 7.3 km,allowing the parafoil to be re-leasedandextractedby theparachute.Theparachuteisdetached fromtheparafoil,afterwhichtheparafoilisinflatedby gradually disreefingindividualsectionsusingaramairsystem[30].The ac-celerationsandvelocitiesduringthisprocesscanbeseeninFig.11. Thefirstpeakdecelerationat2secondscoincideswiththe deploy-mentofthedrogueparachuteat8.1kmaltitude.Thesecondpeak occursbetween7and8seconds andistheresultof parafoil de-ployment.TheparafoildesignandsizingisinspiredbytheMegafly system developed by Airborne Systems [13]. The surface area of the parafoil issized to 836m2,using the desiredmaximum de-scentvelocityandagainamodelofsimpleliftanddragequations asafirstapproximation.

Asixdegreeoffreedom,twelvestateatmosphericflightmodel isdevelopedtosimulatethedynamicbehaviour oftheparafoilin flightandto aidin thedesignof theguidancesystems basedon [27,30,7,31]. This model uses three degrees of freedom for posi-tion and three for orientation. The latter three use Euler angles which contain singularities at certain orientations.It is assumed theparafoilwillflymostly straightandlevelbasedonflight

her-itagedata [5].Therefore useofEuleranglesin themodelis con-sideredacceptable.

The parafoil is controlled using an Air Guidance Unit(AGU),

which contains thecontrol systemhardwareandtwo servosthat control two elevons at the trailing edge with a span of 20.7 m andchordof3.5meach.Theelevonscaneitherbedeployed sym-metricallyto act asbrake,or asymmetricallytoturn theparafoil. The maximumturn performance ofthe control systemislimited to5degreespersecond,topreventheavyoscillationswhichoccur wheneverthisturnrateisexceeded.

The AGU does not require any inner control loops as the parafoil is designed to be inherently stable. Oscillations are present, but verification using models of [31] showed these do notposeaproblemforsafetyorperformance.Therefore,theAGU onlyhasa headingcontrol function,whichitperforms usingtwo different modes: a trajectory mode and an energy management mode.Thetrajectorymodesteerstheparafoiltowardsthepick-up point,immediatelyafterparafoildeployment.Thetrajectorymode controlstheparafoil usingaProportionalandDifferential(PD)

con-troller thatcontrolsthe desiredheading oftheparafoilby means ofelevon deflections.The differentialcontrol isrequiredtodamp out theoscillations thatare aresultofwindgustsaswell asthe payload suspended below the parafoil acting as a pendulum in turning maneuvers.The energymanagement mode isused when theVuABisatthepick-uppointtoloiterinasquarepatternabove theretrievallocation topreparefortheMAR,whichisbetween4 and1.2kmaltitude.

At an altitudeof 4km, thepre-catch phase begins,when the helicopterstartstoaligninfrontoftheVuAB.TheSikorskyCH-53K KingStallionisselectedasasuitablehelicopterforthemission.It has a payload carrying capability of 14500 kg at sea level [29]. The CH-53Kwillinitiatethe catchphase atanaltitudeof1.2km altitude. An InternationalStandardAtmosphere(ISA) model is used to determine that the helicopter can generatethe amount oflift requiredtoperformthecatchat1.2kmaltitude [12].

When the helicopter is aligned in front and above the VuAB andthe velocity vectorsare matched,the couplingmanoeuvre is

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Fig. 12. An overviewofthecatchmechanismusingatandemparafoilconfiguration.

initiated.Thisattachmentisperformedusinganon-clampinghook thatissuspendedfromacableandstabilisedwithadroguebasket. Atthemomentofcatching,thecableallowsahorizontaldistance of2rotordiametersbetweentheVuABandthehelicopterto miti-gatetheeffectsofhelicopterdownwashontheparafoilandcatch mechanism. The hook attaches toa secondary parafoil on top of themain parafoil, alignedwiththe leading edge ofthe latter,as canbeseeninFig.12.Thesecondaryparafoilcarriestheload car-ryingcablethatisattachedtotheVuAB.

Tominimisethecatchingloadsonthehelicopterduringcatch, an attenuation device is used allowing a controlled 8 m expan-sion of the catching cable on the moment of contact with the secondary parafoil. A conservative mass estimation of the catch mechanism mounted on the helicopter is made based on refer-encedata[23],resultinginatotalmassof738.8kg.Thisincludes the entire hoist system: a drum assembly, reeving system, hy-drauliclines,valves,electricsystemcomponents,helicopterdrogue lineanddrogueparachute.Theattenuation[23] reducesthe max-imumcatch loadonthehelicopterfrom180kNto123kNwhich iswithin the helicopters envelope[20]. After the hook is locked on,the



V between theVuABandhelicopterisattenuated,after whichthecableisreeleduntilthepointofattachmentisreached, leavinga linelength of42.21 mbetween thehelicopter andthe VuAB.Uponthisprocessthemainparafoilisdiscardedastheflight dynamicsoftowingtheVuABandthemainparafoilareunstable.

Toensure stable helicopter flight after catch, rotational rates, velocitiesand Eulerangles of the helicopterand VuAB combina-tionare modelledusingan 8-state 3D pendulummodel.Forsafe operation, it is important that the systemdemonstrates stability upon load introduction for the catch, to prevent the pilotsfrom havingtoperformtoodifficultmanoeuvres. Toassessthis perfor-mance,simulations basedon [14] areperformedusing theinitial conditionsgiveninTable6.Thesimulations indicatethat forzero pilotinput,thesystemisstableinroll,butshowsstronginstability aroundthepitchaxis,leadingtodivergingpitchanglesaswellas velocity.Aconstantcontrol surfacedeflectionupon catchis capa-bleofopposing this.It ishoweveradvisabletousean automated controllerforincreasedsafety,andadditionallytodampenoutthe slightoscillationsaroundthepitchangles.

8. Configuration

Thetotalrecoverysystemmass amountsto 2789kg,whichis tobe addedtothefirst stageofthe Ariane6.Internal sourcesat

Table 6

InitialPendulumConditionsforthe8-statehelicoptermodelaftercatchinSection7.

θ0(rad) ˙θ0(rad/s) ¨θ0(rad/s2)

Lateral 0.009 0.018 0

Longitudinal -0.49 -0.030 0

Fig. 13. Cross-section oftheVuABwithfoldedaeroshellandparachute.(For inter-pretationofthecoloursinthefigure(s),thereaderisreferredtothewebversionof thisarticle.)

Airbus indicated that the payload penalty to GTO would realisti-cally be around 0.25 kilograms for each added kilogram of first stagemassforAriane6.Therefore,themodificationswillinducea payloadpenaltyof720kgor16%and6%oftheAriane62and64 GTOpayloadcapacityrespectively[18].Therecoverysystem pack-agingisvisualisedinthecross-sectionoftheVuABinFig.13.The Vulcain 2.1 engine is mounted belowthe green structural cross. Thefuelpipesavoidthecrossanddonotinterferewiththefolded deploymentmechanism.

Four attitude thrusterclusters, each containing 4 thrusters in thesameplanearemountedontheoutsideofthelauncheratthe heightofthestructuralcross.Theseclustersareplacedatanangle of90◦ withrespecttoeachother,tofacilitateintegrationwiththe VuAB.

The parachute is placed opposite to the centre of the folded aeroshelltominimisetheasymmetricmassshiftduetothe recov-ery system. The aeroshellfolding patternis determined usingan assumedminimum foldingradiusof 152mm basedon a ablator thickness of38 mm and a thicknessof 20 mm for the support-ing inflatable toroids [9]. The top of the folded aeroshell has a verticalmarginof100mmwithrespecttothebottomofthe pro-pellant tank whereas the bottomof thefoldedaeroshell rests on the structural cross. The deflated aeroshell follows an accordion foldingpattern. Tofit inside theVuAB, theselayers are wrapped around the longitudinal axis of the launcher. The centre of the aeroshellismountedtotheaeroshelldeploymentmechanismand locatedataradiusof2100mmfromthecentreinfoldedposition. ThisleadstoashiftoftheCentreofGravity(CG) ofthelauncherof 57mmperpendiculartothelongitudinalaxis.

9. Discussion

Thissectionelaboratesonthesignificanceoftheworkaswell asonthe limitationsofthedesign.Firstly, akey designprinciple is to make use of systems with a TRL level that will allow de-velopmenttimewithin4years.HenceallcomponentsoftheVuAB recoverysystemhaveaTRLof5to7.Itimpliesthatthetechnology of those components has been demonstrated and system devel-opment is underway [32]. This is important for the operational

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8 C. Dek et al. / Aerospace Science and Technology 100 (2020) 105778

feasibility of the concept. The Ariane 6, for example, is planned to have the maiden flight late 2020, which is earlier than the moment that the recoverysystemcould be operational afterthe neededdevelopment time.Theintegrationoftherecoverysystem isdesignedsuchthatitrequiresminimalstructuralchangesforthe existinglauncher,onlyrequiringminoradjustmentsinthe attach-mentpointsofthestructuralframe.Thismeansthatdevelopment, testingandintegration oftherecoverysystemcan bedone with-outnecessaryinvasivemeasuresintothedevelopment,production andoperationsprocessoftheAriane6.

The current design is very sensitive to separation velocity. In case of using one boat and one helicopter for the final mid-air catch andlanding, the allowable deviationfrom nominal separa-tion velocity is 0.6%. This is a critical and possibly limiting pa-rameterinthefeasibility ofthedesign.Currently, theVulcain 2.1 engine is not re-ignitable. This is a major driver forthe concept that isdeveloped. Followingcurrent launcher developments,itis expectedthat a majorityof futuregeneration launchers havethe possibilityofre-ignitableengines.Thiswillopenupmanydifferent designoptions,such asthewell-knownboosterlanding.However, itwouldalsoallowforasignificantlylowerrestrictiononthe de-viationofseparationvelocity.

Lastly,asignificantresultisthelowpayloadpenaltyof720 kg. ThisislowcomparedtothepayloadcapacityoftheAriane 6, espe-ciallyconsideringthefactthatthemostexpensivecomponentcan be refurbished and re-used. To assess the cost reduction, a cost breakdownwascreatedthatassessedthecostperlaunchfor vari-ousscenarios.Thisisdoneusingaparametrictopdownapproach using the Transcost model by Dietrich Koelle [17]. This strategy was employed due to the lack of accurate component level cost dataforabottomupapproach.Withthismodel,thedevelopment cost,operationalcostsandproductioncostswereestimated.These weresubsequentlyusedtodeterminetheaveragecostsperlaunch forvariousscenarios inthe lifetimeofAriane 6.This analysis in-corporated a learning factor on the production costs, while the numberoflaunchesper yearandthenumberoftimesanengine canbereusedwerevariedinthesescenariostoevaluatethe sensi-tivitytotheseparameters.Theseindicatedthatforaconservative scenarioof10launchesper yearoverthecourse of20yearsand 4reusesforeachengine,theaveragecostreductionper launchis expectedtobe 15%.

10. Conclusionsandrecommendations

Thedesignofthefirststageheavylaunchvehiclerecovery sys-temfacilitatesseparationoftheVuABandfirststage,deployment anddiscardingoftheaeroshellandparachute,deploymentofthe parafoil, and allows a helicopter to catch the VuAB and land it safelyonavessel.Atlaunch,therecoverysystemwillweigh2789 kgwithapayloadpenaltyof720kg.Thesystemcanbeintegrated intotheexistingdesignofthelauncherandwillnotinterferewith nominal operations of the launcher. Implementing the recovery system can reduce the cost per launch of an Ariane 6 by 15%. Hence,both technical- andeconomic feasibilityaredemonstrated fortheVuABrecoverysystem.GiventheTRLof5-7,adevelopment timeof4yearsmaximumshouldbepossible,allowingforaquick adaptationofthesystem.

Additionalimprovements ontheVuAB recoverysystemdesign canbe realizedthrough furtherinvestigationof everypartofthe designduetothenatureofconceptualdesign.Furtherworkonthe followingconsiderationsisrecommended:

Amore accurate heatflux model during re-entrycould indi-catethatitispossibletouseaninsulatorinsteadofanablator, whichwouldreduce theaeroshell massby 30%. Thiscan

in-creasethere-usabilityandsustainabilityoftheVuABrecovery system.

A more detailed stability and controllability assessment can determinewhetherany(active)control elementsarerequired duringthereentryandaeroshelldiscarding.

Thecurrentparachuteandparafoilmaterials donot allow re-use after contact with ocean water. Improvements made to these materials could allow for re-usability which could re-duce the overall costs and increase the sustainability of the VuABrecoverysystem.

ThepackagingconfigurationcausesaCGoffsetof57mm per-pendiculartothelongitudinalaxisofthelauncher,whichhas to be compensated by a counterweight. Optimisation of the packagingconfigurationmightreducethisCGoffset,removing theneedofacounterweight.

Increasing the space inside VuAB by changing the design of theliquidheliumtankorthestructuralcrosswouldallowthe aeroshell, parachute, and parafoil to be packaged differently, reducingthecomplexityinstructuralattachmentsand deploy-mentmechanisms.

Re-ignition of the Vulcain 2.1 or a similar engine such as Prometheus [11] would allowcontrol of the VuAB after sep-arationandopensupanewarrayofdesignoptionsthatwere discarded inearlierstages ofthecurrentdesign process. The VuAB could land on land, for instance by already giving a thrustboostinspace.

Declarationofcompetinginterest

None.

Acknowledgements

TheauthorswouldliketothankDimitriosZarouchas,at Struc-tures & Materials, and Mario Coppola at Systems Engineering, Aerospace TU Delft, forsupporting the student design team,and Henk Cruijssen, systemengineer Technology & Innovation at Air-busDSNetherlands,forhistechnicalideasandfruitfuldiscussions. Alsowewouldliketothankthreeanonymousreviewersfortheir commentsandhelpinginimprovingoursubmission.

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